This invention relates generally to gas turbine engines, and more particularly to methods and apparatus for assembling fan assemblies for gas turbine engines.
One key factor in aircraft engine design may be the installation and integration of the propulsion system with the aircraft system. For example, the cross-sectional area at the leading edge of the propulsion system, known as the frontal area of the propulsion system, may significantly impact the overall performance of an aircraft system. Specifically, a reduced engine frontal area generally simplifies the installation of the propulsion system into the fuselage or the engine nacelle. Moreover, if the engine frontal area is reduced, then the overall weight of the engine is also reduced.
Within at least some known engines, a fan assembly influences and/or dictates the size of the frontal area. More specifically, known fan assemblies are sized to enable predetermined operational requirements, such as fan inlet radius ratio and/or specific flow to be achieved. The fan inlet radius ratio is generally a mechanical constraint, wherein the airflow output available from the fan assembly is directly linked to the wheel speed of the stage and may be limited by the materials used in fabricating the fan assembly. In contrast, the specific flow of the fan assembly may be limited by aerodynamic constraints.
Other known engines include a “fan-on-blade”, known as a flade, to enable overall aircraft system requirements, such as reduced noise for commercial supersonic aircraft and engine-to-inlet airflow compatibility. In these engines, fan blades are generally coupled to the last stage of the fan assembly due to the increased inlet radius ratio of these downstream stages. Although beneficial, the use of flades may be limited, and more specifically, flades may not be available for use in engines having a reduced engine frontal area, because of the relatively high tip speed of the downstream rotors in engines having a reduced engine frontal area.